Active control of multi-element rotor blade airfoils

ABSTRACT

A multi-element rotor blade includes an individually controllable main element and fixed aerodynamic surface in an aerodynamically efficient location relative to the main element. The main element is controlled to locate the fixed aerodynamic surface in a position to increase lift and/or reduce drag upon the main element at various azimuthal positions during rotation.

The present application is a Continuation-In-Part of U.S. patentapplication Ser. No. 10/147,558, filed 17 May 2002.

This invention was made with government support under CooperativeAgreement: NCC2-9016 for the Variable Geometry Advanced Rotor Technologyprogram awarded by NASA. The government therefore has certain rights inthis invention.

BACKGROUND OF THE INVENTION

The present invention relates to an active multi-element rotor blade,and more particularly to controlling a fixed slat relative throughmovement of a main element.

Multi-element airfoils are in common use on fixed wing aircraft. Suchapplications, however, are either in a fixed configuration or activateat relatively slow rates. In conventional applications, the aerodynamicflow environment is steady or quasi-steady.

Multi-element airfoil application to rotary-wing aircraft hasconcentrated upon the development of fixed elements which attempt toprovide a compromise between achieving an average improvement to rotordisc lift and avoiding an unacceptable increase in drag. Such fixedelements provide numerous design challenges including the aerodynamicrequirements from lower-speed, high angle of attack on the retreatingside of the rotor disc to high speed, low angle of attack operation onthe advancing side of the rotor disc. Current designs for high lift inthe low speed regime suffer from unacceptable drag levels at high speedwhile current designs for low drag in the high-speed regime do not showsufficient benefits of increased lift in the low speed regime.

Accordingly, it is desirable to provide an active multi-element rotorblade airfoil which is configurable to maximize lift performance whileminimizing drag in various flight regimes.

SUMMARY OF THE INVENTION

The present invention provides a multi-element rotor blade having amovable main element and fixed aerodynamic surface, such as a slatpositioned in an aerodynamically efficient location relative to the mainelement. The movable element is controlled via standard blade controlmechanisms including, but not limited to, root actuation control poweredby hydraulic or electrometrical means, flaps or induced structuraltwist. The control means may be operated in accordance withpredetermined parameters which vary according to flight conditions. Inone embodiment, the control means is dynamically responsive to changingflight conditions.

The present invention therefore provides a multi-element rotor bladeairfoil which is configurable to maximize lift performance whileminimizing drag for all flight regimes.

BRIEF DESCRIPTION OF THE DRAWINGS

The various features and advantages of this invention will becomeapparent to those skilled in the art from the following detaileddescription of the currently preferred embodiment. The drawings thataccompany the detailed description can be briefly described as follows:

FIG. 1 is a schematic top view illustrating the aerodynamic environmentof a helicopter main rotor in forward flight;

FIG. 2A is a plan view illustrating a multi-element rotor bladeaccording to the present invention;

FIG. 2B is a plan view illustrating another multi-element rotor bladeaccording to the present invention;

FIG. 2C is a sectional view of the multi-element rotor blade taken alongthe line 2C-2C of FIG. 2B;

FIG. 2D is a sectional view of the multi-element rotor blade taken alongthe line 2D-2D of FIG. 2B;

FIG. 3 is a sectional view of the multi-element rotor blade taken alongthe line 3-3 of FIG. 2;

FIGS. 4A-4F are general schematic views of slat positions relative tothe main element of a multi-element rotor blade according to the presentinvention;

FIG. 5 is a schematic representation of the defined quantities used forthe computed section loads of a multi-element rotor blade;

FIG. 6A is a graphical representation of airloads for multiple slatpositions relative to the main element of a multi-element rotor blade ata Mach number of 0.2;

FIG. 6B is a graphical representation of Steady C_(Lmax) v. Mach numberfor multiple slat positions relative to the main element of amulti-element rotor blade;

FIG. 6C is a graphical representation of steady stall angle v. Machnumber for multiple slat positions relative to the main element of amulti-element rotor blade;

FIG. 7 is a graphical representation of minimum drag coefficients formultiple slat positions relative to the main element of a multi-elementrotor blade;

FIG. 8 is a partial sectional plan view illustrating the elastomericcoupler assemblies for a multi-element rotor blade according to thepresent invention;

FIG. 9A is an expanded view illustrating the elastomeric couplerassemblies of FIG. 8;

FIG. 9B is a sectional view taken along the line 9B-9B of FIG. 9A;

FIG. 9C is a rear view of the elastomeric coupler assemblies of FIG. 9A;

FIG. 10 is a schematic view of the elastomeric bearings illustrating thelayer orientation of the helical and support bearings;

FIG. 11 is an expanded a partial sectional plan view of the root of amulti-element rotor blade illustrating an actuator for the elastomericcoupler assemblies of FIG. 8;

FIG. 12A is a plan view of another multi-element rotor blade having afixed slat according to the present invention;

FIG. 12B is a sectional view of the multi-element rotor blade of FIG.12A according to the present invention having another fixed aerodynamicsurface;

FIG. 12C is a sectional view of another multi-element rotor bladeaccording to the present invention having a fixed aperture aerodynamicsurface;

FIG. 12D is a sectional view of another multi-element rotor bladeaccording to the present invention having another fixed aerodynamicsurface;

FIG. 12E is a sectional view of another multi-element rotor blade havinga flap for individual control of the main element; and

FIG. 12F is a sectional view of another multi-element rotor blade whichis controlled through twisting of the main element.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT

FIG. 1 generally illustrates the aerodynamic environment of a helicopter100 in forward flight having rotor blades of a rotor radius R androtating at an angular speed Ω, while advancing at a flight velocity V.The speed of the airflow over the advancing blade tip is the rotationbased speed ΩR plus the flight velocity V. The airflow over theretreating blade tip is the rotation based speed ΩR minus the flightvelocity V. The airspeed difference between the advancing and theretreating blade is, therefore, relatively large.

The azimuth angle Ψ is an angle measured counterclockwise from the tailof the helicopter. It should be understood that, although Ψ is defined,herein, in reference to a counter clockwise rotating rotor, suchdefinition is for convenience only and should not be consideredlimiting. At an azimuth angle Ψ of 90 degrees, the airspeed of theadvancing blade reaches the maximum of ΩR+V. At an azimuth angle Ψ of270 degrees, the airspeed of the retreating blade reaches the minimum ofΨR−V. The relative airflow at any radial and azimuthal position isobtained by adding a sinusoidal variation of flight speed to therotational speed component, i.e., V(r,Ψ)=Ωr+V sin(Ψ), where r is theradial position and Ψ is the azimuthal position.

FIG. 2A is a plan view showing a multi-element rotor blade 20 having amain element 22 and an active slat 24 movable relative to the mainelement 22. The main element 22 includes a blade root portion 23, acenter portion 26, and a blade tip portion 28. Each portion may define aseparate airfoil section and it should be understood that, although theillustrated embodiment illustrates a particular design, other rotorblades will benefit from the present invention.

The blade root portion 23 is attached to a rotor head (not shown) forrotating the rotor blade 20 about an axis of rotation A. The mainelement 22 defines a leading edge 22 a and a trailing edge 22 b, whichare generally parallel to each other. The distance between the leadingedge 22 a and the trailing edge 22 b defines a main element chord lengthCm. It should be understood that any rotor blade design will benefitfrom the present invention.

The slat 24 is mounted parallel to the leading edge 22 a and is movablerelative thereto by an actuator assembly (illustrated schematically at30 in FIG. 2A and FIG. 11) or the like about elastomeric couplerassemblies (illustrated schematically at 34). The slat 24 defines aleading edge 24 a and a trailing edge 24 b, which are generally parallelto each other. A distance between the leading edge 24 a and the trailingedge 24 b defines a slat chord length Cs.

It should be understood that various actuators and slat mountingarrangements will benefit from the present invention. Actuators such asmechanical, electrical, pneumatic, piezoceramic, hydraulic and the likewill also benefit from the present invention. It should also beunderstood that, although the present invention is described with regardto a multi-element airfoil on a main rotor, many other movable membersincluding airfoil and non-airfoil members will benefit from the presentinvention. Moreover, other coupling applications will also benefit fromthe present invention such as, for example, hinges for doors or thelike.

Additionally or alternatively, an electronic controller 31 operates anactuator 30, which moves the slat 24. It should be further understoodthat although the slat in the disclosed embodiment is illustrated alonga particular longitudinal length of the main element 22 other lengthsand locations for continuous or partial length slats will benefit fromthe present invention.

Referring to FIG. 2B, another blade design includes an active slat 24′,which preferably spans from the 75% R to 90% R length. A plurality ofelastomeric coupler assemblies 34′, upon which the slat 24′ moves, aremounted to the main element 22′ to support rotation and translation ofthe slat 24′ about a virtual hinge point Vh (FIG. 2D). FIG. 2Billustrates another location for actuator 30′ mounted inboard slat 24′adjacent the root (FIG. 11).

Referring to FIG. 3, a sectional view of the rotor blade 20 isillustrated. The leading edge 22 a of the main element 22 defines anorigin of coordinates (x, y)=(0, 0). The base slat geometry and positionpreferably provides acceptable (relative to a conventional singleelement airfoil) maximum coefficient of lift C_(Lmax) at moderate Machnumber (retreating blade) while minimizing drag C_(D) at higher Machnumber (advancing blade). Applicant has determined that the slat chordCs is preferably rotated nose down at an angle (βslat) with respect tothe main element chord Cm to achieve desired base position results. Inone preferred embodiment, in which the main element chord Cm is 24inches and the leading edge 22 a defines the origin of coordinates, theslat leading edge 24 a is at (x, y)=(−1.892″, +0.642″), the slattrailing edge is at (+1.677″, +1.943″) and βslat=20 degrees nose down.The slat quarter chord is at (−0.983″, +0.973″) and the base positionrotation point 32 (shown in FIG. 4A) is at (0.5″, 1.0″.) It should beunderstood, however, that other multi-element rotor blades will benefitfrom the present invention.

Referring to FIGS. 4A-4F, the slat 24 preferably rotates and translatesrelative to the main element 22 from a base position (FIG. 4A). The slat24 preferably rotates about the base position rotation point 32 andtranslates relative thereto. From the base position rotation point 32,the slat 24 also translates along a translation line t (FIGS. 4D-4F)defined down and away from the main element 22. Preferably, thetranslation line is orientated to provide slat 24 translationapproximately 45 degrees down and away from the main element 22.Positive rotation is slat nose up.

Rotation and translation of the slat 24 may also be accomplished bylocating the rotation point 32 outside the contour of the slat 24 andmain element 22, as depicted by the virtual hinge point Vh in FIG. 2D.When extended along the span of the slat 24, the virtual hinge point Vhdefines an axis H (FIG. 10) generally parallel to both the main element22 and the slat 24. That is, by locating the virtual hinge point Vhoutside the contour of the main element 22′ opposite the slat 24′,rotation and translation is obtained, similar to that shown in FIGS.4A-F. It should be understood that other positions will also benefitfrom the present invention. Furthermore, the slat 24′ is continuouslymovable through an infinite number of positions between at least any twoof the representative positions.

The base position (FIG. 4A) provides a compromise between minimumcoefficient of drag C_(D) and maximum coefficient of lift C_(Lmax).Position 2 (FIG. 4B) rotates the slat −5° (nose down) about x, y=(0.5″,1.0″). Position 3 (FIG. 4C) rotates the slat +5° (nose up) about x,y=(0.5″, 1.0″). Position 4 (FIG. 4D) provides a −5° rotation with atranslation of x, y=(−0.25″, −0.25″). Position 5 (FIG. 4E) combines a−2.5° rotation about x, y=(0.5″, 1.0″) with a translation of x,y=(−0.25″, −0.25″) from x, y =(0.5″, 1.0″). Position 6 (FIG. 4F)combines a −7.5° rotation with a translation of x, y=(−0.45″, −0.45″)from x, y=(0.5″, 1.0″) to maximize C_(Lmax) capability. It should beunderstood that, although discrete positions are disclosed in theillustrated embodiment, other positions will benefit from the presentinvention. In other words, the slat 24 is continuously movable throughan infinite number of positions between maximum positions defined at,preferably, Ψ=90° azimuth position and Ψ=270° azimuth position. Themaximum positions are preferably limited by mechanical stops or thelike.

Referring to FIG. 5, defined quantities used for the computed sectionloads are illustrated. MAIN refers to the loads on the main elementonly, normalized by the main element chord, with the pitching moment atthe main element 1/4 chord. SLAT refers to the loads on the slat only,normalized by the slat chord and using the slat 1/4 chord as the momentcenter. TOTC is total airfoil load, now normalized by the total chord(C_(TOTC)) and with the moment center at the total airfoil 1/4 chord(x_(1/4TOTC)). βslat is the angle of the slat chord with respect to themain element chord.

Locations x and y are referenced to the main element leading edge andparallel and normal to the main element chord line. x_(1/4 TOTC) is the1/4 chord position of the complete airfoil, and (x_(Slat 1/4),Y_(Slat 1/4)) is the location of the slat 1/4 chord. Geometricalparameters for the positions are shown in the Table below. Configurationβ_(Slat) C_(TOTC) x_(1/4 TOTC) x_(Slat 1/4) y_(Slat 1/4) Main Only N/A24.00″ 6.00″ N/A N/A Slat 1 −20° 25.89″ 4.58″ −0.98″ +0.97″ Slat 2 −25°25.89″ 4.58″ −0.98″ +0.84″ Slat 3 −15° 25.89″ 4.58″ −0.98″ +1.10″ Slat 4−25° 26.14″ 4.40″ −1.23″ +0.59″ Slat 5 −22.5° 26.14″ 4.40″ −1.23″ +0.66″Slat 6 −27.5° 26.34″ 4.25″ −1.42″ +0.33″

Referring to FIG. 6A, airloads for the various configurations at a Machnumber of 0.2 are graphed. Mach 0.2 is an example of a retreating bladespeed. Slat position 6 (FIG. 4F) consistently provides the highestcombined C_(Lmax) (C_(LTOTC)) and the latest stall. Steady C_(Lmax) v.Mach number is graphically depicted in FIGS. 6B, and steady stall anglev. Mach number is graphically depicted in FIG. 6C. Slat position 6 istherefore the preferred position for the retreating side of the rotordisc.

Referring to FIG. 7A, minimum drag coefficients for each slat positionare illustrated. The minimum drag values for the main element only, mainelement with slat position 1, and position 3 are relatively close toeach other, rising rapidly only above M=0.75. Position 5 also hassimilar minimum drag values at a lower Mach number, but exhibits rapiddrag rise at M=0.70. Position 3, the only positive rotation slatposition, provides the lowest drag at higher Mach numbers relevant foran advancing blade. Slat position 3 is therefore preferred for theadvancing side of the rotor disc

For the advancing blade, because the airspeed thereof is significantlygreater than the retreating blade, applicant has determined that apositive rotation from the base position (position 3; FIG. 4C and FIG.7) minimizes drag at low angles of attack, providing a coefficient ofdrag C_(D) lower than a conventional single element rotor blade. For theretreating blade, since the airspeed thereof is significantly lower thanthe advancing blade, Applicant has determined that negative rotation andtranslation (position 6; FIG. 4F and FIG. 6B) maximizes the coefficientof lift C_(Lmax). Slat position 6 provides an increase in the steadystate coefficient of lift by 0.3 to 0.45 and also an improved averagelift capability for unsteady motions.

Actuation of the slat can be prescribed for a given aircraft, prescribedin response to a given aircraft at a given flight condition, or activelycontrolled in that real time sensors, which acquire data and via acomputer processing algorithm within controller 31, demand slat positionto optimize a defined performance objective. Prescribed motion is openloop; there is no feedback. In other words, a prescribed motion scheduleis defined and the slat moves accordingly (although a differentprescribed motion may be provided for different flight conditions orloading conditions, etc.—but these are relatively slowly changingconditions).

In one control schedule, the actuator 30 (FIG. 2A) controls the movementof the slat 24 in a prescribed manner in accordance with azimuth angleΨ. The actuator 30 may be controlled by a cam arrangement or the like,which drives the slat in a predetermined pattern relative to the azimuthangle Ψ. Preferably, the slat 24 is driven between slat position 6 andslat position 3 in a sinusoidal wave pattern, which has a maximum noseup position near Ψ=90 degrees and a maximum nose down position nearΨ=270 degrees. That is, the slat 24 is at a low drag position (position3; FIG. 4C) at Ψ=90 degrees, at high lift position (position 6; FIG. 4F)at Ψ=270 degrees and at a base position (approximately position 1; FIG.4A) at Ψ=0, 180, and 360 degrees. It should be understood that it maynot be physically possible to pass exactly through the 3 positionsillustrated by FIGS. 4C, 4F and 4A (positions 1, 3 and 6) with a singlepivot axis H. With single pivot axis H located to achieve positions 1and 6, the slat will move through a base position which may differ fromthat which is illustrated as position 1 (FIG. 4A).

Prescribed motion is based on providing a minimum drag configuration inthe high speed region (advancing blade), where drag is more critical foroverall rotor performance, while also providing a maximum liftconfiguration in the lower speed region (retreating blade), where liftis more critical for overall rotor performance. Such a prescribed motionmay be defined as, for example only, a once per revolution (1P)sinusoidal motion. A 1P function can be defined as: y=A*sin(1*Ψ), where“A” is the amplitude and “Ψ” is the azimuthal angle. The 1P functionprovides a smooth continuous curve that starts at zero at Ψ=0 (bladeover the tail), grows to an amplitude of “A” at Ψ=90 degrees (FIG. 1;blade pointing directly to the right side—maximum advancing speed),diminishes back to zero at Ψ=180 degrees (blade over the aircraft nose)and then further diminishes to an amplitude of minus “A” at Ψ=270degrees (FIG. 1; blade pointing directly to the left—minimum speed,called the retreating blade side), and then finally grows back to zeroat Ψ=360 degrees which equals Ψ=0 degrees (blade over tail). A 2P (ortwice per rev) function: y=A*sin(2*Ψ) provides the same IP curve shape,only it completes a complete cycle in ½ the azimuth—between Ψ=0 andΨ=180. So for a single revolution, 2 cycles are achieved—which means twomaximum points of amplitude “A,” (at 45 degrees and at 225 degrees) andtwo minimums of amplitude, “−A,” (at 135 degrees and at 315 degrees). Ina like manner, a 3P function has 3 cycles per rotor revolution and a 4Pfunction 4 cycles per revolution, and so forth. Superimposing and phaseshifting of control harmonics can be used to alter the prescribed motionin response to flight condition to provide an optimal response ordesired characteristics. That is, 1P prescribed motion may be preferredfor one flight condition, while 1P and/or XP may be preferred foranother flight condition.

A controller determines the optimal fanction to alter the prescribedmotion schedule being enacted. In this case, however, the rate at whichan aircraft changes flight condition (such as speed or altitude) is farbelow the rotational speed of the rotor, so this more rigid (functionmapping) scenario is defined as prescribed control. Further, forprescribed motion a somewhat arbitrary distinction may be definedbetween “simple” 1P motion (the slat moves to low drag on the advancingside and high lift on the retreating side), and more complex, higherharmonic motion which essentially becomes an arbitrary motion (which canbe defined as a sum of harmonics), but is still prescribed withoutrequiring a closed loop feedback system.

Changing of prescribed motions, due to changing from one flightcondition to another, has aspects of an active control system. That is,control logic selects one of several predefined prescribed motions basedon sensor measurements, such as flight speed and altitude. The“prescribed control” strategy differs from active control, wherebysensor feedback (at rates similar to or higher than rotor speed) is usedto change slat motion up to a per rev basis.

Prescribed motion may additionally, or in the alternative, include otherprescribed motion functions, such as moving the maximum slat nose upand/or slat nose down positions to other points in the azimuth, and/orsustaining a given deflection for some period, or other motions thatoptimize defined performance objectives. These additional alternativemotion functions are defined as a sum of sinusoidal motions of differentfrequencies and phase angles, i.e. how rapidly the sinusoidal motionoccurs and the relative starting point of the motion around the azimuth.The “phase” of the function changes the starting point of the motionaway from Ψ=0 degrees to any other point, and the sharper a motion gets,the more harmonic functions are required to the limit in which a stepchange function requires an infinite number of harmonic functions.

For active or adaptive control, sensor data is acquired in real time andas a result of this data, the slat motion is controlled to provide anoptimized motion based on predetermined objectives. Slat motioncommanded by closed loop control has an arbitrary waveform (within theconstraints of the controller basis functions) in addition oralternatively to prescribed motion. Active control tailors the slatmovement in real time, as sensor data is acquired and processed throughdefined algorithms, to meet a defined objective function. This objectivefunction can be made up of multiple performance objectives and tailoredfor different modes of operation, e.g. a high performance mode, a lownoise mode, etc., and for different flight conditions and/orconfigurations, e.g., hover, forward flight, air-to-air engagement, etc.As stated with regard to prescribed motion, the active control motionsare defined in the time context of the order of the blade rotation, i.e.the processing and sensor feedback are at a rate similar to the rotorrotation rate. Implementation of active control may additionally oralternatively include prescribed motion functions in response to baseconditions, e.g., forward flight, hover, etc.

Referring to FIG. 8, an expanded view of the multi-element rotor blade20 illustrates the elastomeric coupler assemblies 34 a, 34 b thatmovably support the active slat 24 relative to the main element 22. Itshould be understood that although described with regard to movement ofa multi-element rotor blade virtually any coupling between linear orrotary degrees of freedom including airfoil and non-airfoil members willbenefit from the present invention.

Preferably, an inner elastomeric coupler assembly 34 a and an outerelastomeric coupler assembly 34 b support the slat 24 therebetween. Anactuator rod 36 extends within the main element 22, from the blade rootportion 23 (FIG. 11) to the inner elastomeric coupler assembly 34 a, toactuate the slat 24. Spanwise actuation is particularly desirable,because of multi-element airfoil geometric constraints, structuralconstraints, and component mounting considerations.

The actuator rod 36 is preferably a tension rod, which is only “pulled”by the actuator 30. Centrifugal force operates to drive the slat 24 to afirst position and the actuator rod operates in tension to pull upon theinner elastomeric coupler assembly 34 a to drive the slat 24 inopposition to the centrifugal force to a second position. It should beunderstood that other actuators, which provide other inputs such as arotational input, will also benefit from the present invention.

Referring to FIG. 9A, an expanded view of the inner and outerelastomeric coupler assemblies 34 a, 34 b is illustrated. Eachelastomeric coupler assembly 34 a, 34 b includes a grounding member 38and an active member 40. The grounding member 38 is fixed to the mainelement 22 by fasteners 42, such as bolts or the like, which extend intothe spar 44 of the main element 22. The grounding member 38 ispreferably recessed within the leading edge 22 a (also illustrated inFIG. 8) toward the spar 44 of the main element 22 such that slat 24 maybe at least partially embedded in the planform.

The active member 40 supports the slat 24. Fasteners 46, such as boltsor the like, secure the active member 40 to a spanwise end of the slat24 (FIG. 9B). The active member 40 is movably mounted to the groundingmember 38 through a helical elastomeric bearing 46, and a first andsecond elastomeric support bearing 48 a, 48 b (also illustrated in FIG.9C).

The helical elastomeric bearing 46, and the first and second elastomericsupport bearings 48 a, 48 b include a plurality of layers of sheardeformable elastomeric material layers 50 separated by helical shimlayers 52 formed of high-stiffness constraining material (FIG. 10) suchas composite or metallic layers. It should be understood, however, thatvarious materials of differing rigidity will also benefit from thepresent invention. The helical elastomeric bearing 46 operates as acoupler between the active member 40 and the grounding member 38. Underaction of centrifugal force, the shear deformable elastomeric materiallayers 50 shear within the constraints of the helical shim layers 52(FIG. 10). The helical shims 52 guide the elastomer shear deformationsuch that the displacement trajectory of the active member 40 relativeto the grounding member 38 is a predefined coupled spanwise-translationand rotation.

The elastomeric bearings 46, 48 a, 48 b preferably define an arcuate orcupped shape within the plane of the blade section (FIG. 9B) having afocus at the desired virtual hinge point Vh (FIGS. 2D and 9B). Theelastomeric bearings 46, 48 a, 48 b thereby support rotation of the slat24 about the virtual hinge axis H. As the virtual hinge axis H is notcoincident with the slat 1/4 chord, slat rotation between two points onan arc about the hinge axis H provides simultaneous rotation of the slatchord and translation of the slat 1/4 chord. The arc defined about hingeaxis H provides single arc motion preferably between position 3 (FIG.4C) and position 6 (FIG. 4F). In particular, the slat rotates about theaxis H and the motion is completely described by a single arc angleparameter.

In order to achieve slat motion from the desired low drag position(Position 3, FIG. 4C) to the high lift position (Position 6, FIG. 4F),for the current blade geometry, the virtual hinge point is locatedoutside the contour of the main blade section. Depending on the locationof Vh, single arc motion between any two other points can be achievedwithin elastomeric deformation limits. It should be understood that toachieve motion through other predefined positions the virtual hingepoint may be located in another position relative the contours.

The grounding member 38 defines a support ramp 54 upon which the helicalelastomeric bearing 46 acts against. The support ramp 54 is angled awayfrom the axis H at an acute angle in substantially the same plane of theleading edge 22 a and the trailing edge 22 b (FIG. 8). That is, theacute angle is formed by the outboard face of the grounding member andthe axis H. The support ramp 54 of the inner elastomeric couplerassembly 34 a is oriented the same way as the support ramp 54 of theouter elastomeric coupler assembly 34 b. That is, the angle definedbetween axis H and the support ramp 54 of the inner elastomeric couplerassembly 34 a faces toward the slat 24, while the same angle between theaxis H and support ramp 54 of the outer elastomeric coupler assembly 34b faces away from the slat 24 (FIG. 8).

The support ramp 54 of the inner elastomeric coupler assembly 34 a andthe outer elastomeric coupler assembly 34 b are preferably concave (FIG.9C). A mating ramp 55 of the active member 40 of the inner elastomericcoupler assembly 34 a and the outer elastomeric coupler assembly 34 b,which faces the support ramp 54, are preferably convex (FIG. 9C). Thehelical elastomeric bearing 46 is located between the concave supportramp 54 and the convex mating ramp 55. The purpose of the curvature ofthe support and mating ramp, elastomeric layers, and high stiffness shimlayers, as seen in the plane of the blade (FIG. 9C), is to stabilize theelastomer and thereby stabilize the slat 24.

The elastomeric support bearings 48 a, 48 b provide motion by sheardeformation of the elastomeric layers and carry radial loads incompression against fixed upper and lower support caps 56, 60. Thegrounding member 38 defines a fixed upper support cap 56 for the firstelastomeric bearing 48 a, and a fixed lower support cap 60 for thesecond elastomeric bearing 48 b (FIG. 9C). The active member 40 isthereby trapped by the first elastomeric bearing 48 a acting upon theupper cap and the second elastomeric bearing 48 a acting upon the lowercap 60. The first elastomeric support bearing 48 a operates to carry thenominal upward lift upon the slat while the second elastomeric supportbearing 48 b operates to support the slat in the event that there is adownload on the slat. Download may occur when the outboard portion ofthe main element operates at a negative angle of attack, i.e., when theoutboard portion of the blade “digs-in”.

Referring to FIG. 10, the helical elastomeric bearing 46, and theelastomeric support bearings 48 a, 48 b are schematically illustratedrelative to the virtual hinge point Vh (axis H), for the outboardbearing assembly. The elastomeric support bearings 48 a, 48 b are cuspedrelative to the virtual hinge point such that they define an arc in theplane of the blade section, which has a focus point on the virtual hingeaxis H. The layers of the elastomeric support bearings 48 a, 48 b arearranged substantially parallel to the virtual hinge axis H. That is, aplane which is perpendicular to the virtual hinge axis H would passthrough all the layers which define the elastomeric support bearings 48a, 48 b.

The first and second elastomeric support bearings 48 a, 48 b provide afirst degree of freedom axially along the virtual hinge axis H. Thefirst and second elastomeric support bearings 48 a, 48 b provide asecond degree of freedom about the virtual hinge axis H, which providesfor the rotation of the slat 24 (FIG. 9B) about the virtual hinge axisH. These two degrees of freedom are independent of each other. Theradial load capacity of the elastomeric support bearings 48 a, 48 bdepends upon the caps 58,60 (FIG. 9C), the composition of the layers,and the extent width of the bearing.

The helical elastomeric bearing 46 is cusped such that, as with theelastomeric support bearings 48 a, 48 b, the helical elastomeric bearing46 defines an arc (in the plane of the blade section), which has a focalpoint generally along axis H. The helical elastomeric bearing 46 islayered such that it defines a section of a circular helix, whichencircles the virtual hinge axis H. Preferably, the helix issufficiently spread out spanwise along a +50 degree helix angle, inwhich adjacent layers are not from one helix locus. That is, theadjacent layers are respective segments of identical spanwise stackedhelix loci. Further, as the helical elastomeric bearing 46 is locatedbetween the concave support ramp 54 and the convex mating ramp 55, thehelical elastomeric bearing 46 is essentially double curved.

The helical elastomeric bearing 46 is layered such that it (as opposedto the elastomeric support bearings 48 a, 48 b) defines a section of acircular helix. The helical elastomeric bearing 46 and support ramp 54are angled in the same direction as the helix angle. The layers of thehelical elastomeric bearing 46 are arranged substantially about thevirtual hinge axis H. That is, a plane which is perpendicular to thevirtual hinge axis H would not pass through all the layers of thehelical elastomeric bearing 46. The helical elastomeric bearing 46converts a linear input parallel to the virtual hinge axis H into arotary output to rotate and translate the slat.

Referring to FIG. 11, the actuator rod 36 is driven by the actuatorassembly 30, which preferably includes an actuator 62, such as ahydraulic, pneumatic, electric, mechanical, electromagnetic,piezoceramic actuator, or the like. It should be understood that theactuator 62 may additionally or in the alternative be a mechanicallylinkage which receives control inputs through a fixed amplitudeswashplate or the like. The actuator 62 is preferably located adjacentthe blade root portion 23 and drives the actuator rod 36 through a crank64 to provide mechanical advantage thereto. The crank 64 also reversesthe motion of the actuator such that an extension actuator can bemounted with a fixed end inboard. The actuator rod 36 is preferably atension rod which extends within the main element 22, from the bladeroot portion 23 to the inner elastomeric coupler assembly 34 a, toactuate the slat 24 (FIG. 8).

In operation, centrifugal force operates to slide the slat 24 outboardtoward the blade tip. The elastomeric support bearings 48 a, 48 b of theelastomeric coupler assemblies 34 provide minimal shear resistance tothis sliding movement. In the helical elastomeric bearing 46, however,the spanwise outboard sliding motion of the slat 24 acts as an input tothe intrinsic helical coupling of the elastomeric bearing 46, resultingin an output rotation of the slat 24. The active member 40 isconstrained to move along the helical arc relative to the groundingmember 38 by means of incremental shear of the elastomeric layers 50between the support ramp 54, the respective shim layers 52 and themating ramp 55. The centrifugal force is reacted through the tension rod36.

Outboard sliding of the slat 24 is accommodated by the active member 38and the attached slat 24 moving elastomerically along the helical arc ofthe helical elastomeric bearing 46 such that the slat 24 rotates nosedown. Preferably, maximum travel of the slat is mechanically limited tothe fully deployed position 6 (FIG. 4F and FIG. 6) to maximize thecoefficient of lift C_(Lmax). This arrangement negates the need for asafety interlock, as in the event of a hardware/software failure theactuator 62 need only be vented and the slat will achieve its fullydeployed position. Moreover, as centrifugal force operates to drive theslat to a deployed condition, the slat requires powered activation foronly one direction. Nose up pitching moment upon the slat tends torotate the slat up and the helical coupling will move the slat inboardpartially unloading the actuator.

Applicant has determined that a helical elastomeric bearing with a 58degree helix angle and ±0.5 inches of spanwise travel achieves a 10degree peak to peak slat rotation about the virtual hinge point Vh.

To retract the slat 24, the actuator 62 operates to place the actuatorrod 36 under tension to pull upon the active member 40 of the innerelastomeric coupler assembly 34 a. Retraction of the active member 40relative to the grounding member 38 retracts the slat 24 in oppositionto centrifugal force. Preferably, the controller 31 (FIG. 2A) controlsthe actuator 62 to drive the slat to a desired position under prescribedor active control as described above.

Referring to FIG. 12A, a fixed element such as a fixed slat 24 f isfixed to the leading edge 22 a′ of the rotor blade 20′ and remainsstationary relative a main element 22′ while the rotor blade 20′ rotatesabout the axis of rotation A (also illustrated in FIG. 12B). Although afixed slat 24 f is disclosed in the illustrated embodiment other fixedelements that are located on the blade 20′ such as a slot s (FIG. 12C),a wire W (FIG. 12D), and/or other fixed protrusion mounted parallel tothe leading edge 22 a′ will benefit from the present invention. Eachfixed element provides for aerodynamic modifications to the rotor blade20′ in response to individual blade control.

Rotor blade 20′ includes a flight condition sensor 70 for determiningflight characteristics of the rotor blade during rotation about axis A.A controller 72 communicates with the sensor 70 to sense a position ofthe blade 20′ and provide individual blade control in response toazimuthal position. That is, each blade 20′ is individually positionedas it rotates about axis A by changing the aerodynamic characteristicsof the main element 22′. Preferably, each rotor blade 20′ is controlledto maximize the aerodynamic performance of the fixed element incombination with the main element 22′ of the blade 20′ to provide areduction in drag on the main element 22′ when the aircraft is operatingat higher ranges of speed and/or to provide an increase in lift whenoperating in a lower range of speed. That is, by individuallycontrolling the main element 22′, the fixed element 24 f is positionedto improve the aerodynamics of the entire rotor blade 20′.

Referring to FIG. 12B, the rotor blade 20′ is preferably controlledthrough individual pitch control about a pitch rotation axis 74 relativeto azimuthal position. That is, rotor blade 20″ is controlledindependently of a plurality of rotor blades which are all mounted to acommon rotor head (FIG. 1). It should be understood, that variousactuation methods such as independent root pitch actuation control ofeach individual blade, deployment of a flap 76 (FIG. 12E) to manipulatethe rotor blade 20′ about the pitch rotation axis 74, and/or or inducinga structural twist of the rotor blade 20′ (FIG. 12F) will benefit fromthe present invention.

The controller is responsive to a prescribed motion schedule relative toazimuthal position and in response to changing flight conditions thatmay or may not be alterable according to flight conditions as describeabove. It should be understood that various combinations of fixedelements along with other individual blade control methodologies willbenefit from the present invention.

The foregoing description is exemplary rather than defined by thelimitations within. Many modifications and variations of the presentinvention are possible in light of the above teachings. The preferredembodiments of this invention have been disclosed, however, one ofordinary skill in the art would recognize that certain modificationswould come within the scope of this invention. It is, therefore, to beunderstood that within the scope of the appended claims, the inventionmay be practiced otherwise than as specifically described. For thatreason the following claims should be studied to determine the truescope and content of this invention.

1. A multi-element rotor blade for an aircraft comprising: a mainelement rotatable about an axis of rotation; an second aerodynamicsurface fixed along a leading edge said main element; a control forpitching said main element at a rate greater than once per revolutionsuch that said second aerodynamic surface is positioned to enhance theaerodynamic performance of said main element relative to azimuthalposition.
 2. The multi-element rotor blade as recited in claim 1,wherein said control comprises a root actuator to pitch said mainelement.
 3. The multi-clement rotor blade as recited in claim 1, whereinsaid control comprises a flap to pitch said main element.
 4. Themulti-element rotor blade as recited in claim 1, wherein said controlcomprises an actuator to twist said main element along a longitudinalaxis.
 5. The multi-element rotor blade as recited in claim 1, whereinsaid control operates to move said main clement at a rate greater thanonce per revolution when said main element is disposed among a pluralityof other main elements.
 6. A multi-element rotor blade for an aircraftcomprising: a main element rotatable about an axis of rotation; anaperture through said main element; a control for moving said mainclement at a rate greater than once per revolution such that saidaperture is positioned to enhance the aerodynamic performance of saidmain element relative to azimuthal position.
 7. The multi-element rotorblade as recited in claim 6, wherein said control comprises a rootactuator to pitch said main element.
 8. The multi-element rotor blade asrecited in claim 6, wherein said control comprises a flap to pitch saidmain element.
 9. The multi-element rotor blade as recited in claim 6,wherein said control comprises an actuator to twist said main elementalong a longitudinal axis.
 10. A method of controlling a multi-elementrotor blade comprising a second aerodynamic surface fixed along aleading edge of a main element among other multi-element rotor bladeswhich rotate about a common axis of rotation, said method comprising thesteps of: (1) controlling a pitch position of said main element at arate greater than once per revolution such that said aerodynamic surfaceis positioned to enhance performance of said main clement.
 11. A methodas recited in claim 10, further comprising the step of: controlling thepitch position of the main element at a rate greater than once perrevolution such that the aerodynamic surface is positioned to decreasedrag of the main element.
 12. A method as recited in claim 10, furthercomprising the step of. controlling the pitch position of the mainelement at a rate greater than once per revolution such that theaerodynamic surface is positioned to increase lift of the main element.13. A method as recited in claim 10, further comprising the step of:pitching the main element at a root of the main element at a rategreater than once per revolution.
 14. A method as recited in claim 10,further comprising the step of: twisting the main element at a rategreater than once per revolution.
 15. A method as recited in claim 10,further comprising the step of: extending an aerodynamic surface fromthe main element to pitch the main element at a rate greater than onceper revolution.
 16. A method as recited in claim 10, further comprisingthe step of: controlling the main element in accordance with aprescribed motion schedule.
 17. A method as recited in claim 16, furthercomprising the step of: modifying said prescribed motion schedule inresponse to a flight condition.
 18. A multi-element rotor blade for anaircraft comprising: a main element rotatable about an axis of rotation;an second aerodynamic surface fixed to said main clement; a control forpitching said main element at a rate greater than once per revolutionsuch that said second aerodynamic surface is positioned to enhance theaerodynamic performance of said main element relative to azimuthalposition, said control comprising an actuator to twist said main clementalong a longitudinal axis.
 19. A method of controlling a multi-elementrotor blade comprising an aerodynamic surface fixed to a main elementamong other multi-element rotor blades which rotate about a common axisof rotation, said method comprising the steps of: (1) controlling aposition of said main element at a rate greater than once per revolutionsuch that said aerodynamic surface is positioned to decrease drag of themain element.
 20. A method of controlling a multi-element rotor bladecomprising an aerodynamic surface fixed to a main element among othermulti-element rotor blades which rotate about a common axis of rotation,said method comprising the steps of: (1) pitching the main element at aroot of the main element at a rate greater than once per revolution suchthat said aerodynamic surface is positioned to enhance performance ofsaid main element.
 21. A method of controlling a multi-element rotorblade comprising an aerodynamic surface fixed to a main element amongother multi-element rotor blades which rotate about a common axis ofrotation, said method comprising the steps of: (1) twisting a positionof the main element at a rate greater than once per revolution such thatsaid acrodynamic surface is positioned to enhance performance of saidmain element.
 22. A method of controlling a multi-element rotor bladecomprising an aerodynamic surface fixed to a main element among othermulti-element rotor blades which rotate about a common axis of rotation,said method comprising the steps of: (1) controlling a position of saidmain element at a rate greater than once per revolution in accordancewith a prescribed motion schedule such that said aerodynamic surface ispositioned to enhance performance of said main element; and (2)modifying said prescribed motion schedule in response to a flightcondition
 23. The multi-element rotor blade as recited in claim 1,wherein said second aerodynamic surface comprises a fixed leading slat.24. The multi-element rotor blade as recited in claim 1, wherein saidsecond aerodynamic surface comprises a fixed protrusion.
 25. Themulti-element rotor blade as recited in claim 1, wherein said secondaerodynamic surface comprises a slot.